Chapter 3: Transonic Wind Tunnel Development (1940 -1950)
[84] The idea of using the local region of transonic and supersonic flow that develops on wings at high subsonic speeds as a medium in which useful testing of small aerodynamic models could be accomplished did not occur to high-speed wind tunnel researchers for a simple reason: their airfoil and wing models were generally so small that there was no practical possibility of such an approach. However, starting with the Brewster XF2A-2 airplane dive tests of 1940 (ref. 101) the Flight Division had seen these local transonic flow fields develop on wing sections 10 to 20 times larger than the small wind-tunnel airfoil models. Noting the absence of any constriction effects due to tunnel walls, R. R. Gilruth proposed in 1944 that these aircraft wing flow fields be utilized for transonic testing of small models (ref. 47).
The first reaction of our high-speed wind-tunnel group was quite negative. With our 15-year background of effort at generating uniform flows for valid testing we pointed to the many obvious problems of the wing-flow technique-the flow-field nonuniformities both chord-wise and normal-to-chord, the wing boundary layer, the problem of wing shock passage over the test model, interference due to clearance between model and wall, and the very low test Reynolds numbers which were well below those of our smallest wind tunnel models. Gilruth persisted, however, arguing that any transonic data would be preferable to none. When his results (ref. 47) became available, showing for the first time continuous plots of wing lift, drag, and moment data through Mach 1 and up to about 1.3 and trends which appeared to conform to expectations, we were impressed.
The rather obvious thought that the wing-flow scheme could be applied by mounting a large wing section in one of our high-speed tunnels, equally well or better than on a diving airplane, must have occurred to many in 1945. There was no immediate rush to exploit the idea, however. Many researchers, perhaps a majority, still found the scheme so fraught with problems and impurities as to be unworthy of [85] adoption, and this view tended to prevail in our 16-foot-tunnel group. Nonetheless, early in 1946 we decided to make a quick preliminary check of what might be done by investigating a large-chord airfoil spanning the 16-foot tunnel test section. G. Heiser reported the results at the June 3, 1946, meeting of the Langley General Aerodynamics Committee. He had found that the large airfoil absorbed so much power that the maximum local Mach number reached at full tunnel power was only about 1.0, considerably less than we had hoped for. Heiser estimated that much better performance could be obtained by mounting a short section of the airfoil directly on the floor of the tunnel and fairing it into the wall. Obviously a tunnel boundary-layer removal system would also be required. This would have involved more cost and effort than the idea was worth, in our opinion, and Heiser told the committee we were planning no further work at 16-foot. This investigation is believed to have been the first NACA attempt to define and develop what later came to be called the "bump." Lockheed (ref. 102), Ames (ref. 103), and Langley (ref. 104), started subsequent successful developments of the bump in 1946. It was used extensively in the Ames 16-foot high-speed tunnel and the Langley high-speed 7 x 10-foot tunnel, largely replacing the aircraft wing-flow work in the period before the large slotted tunnels became fully operational. The bump programs naturally disappeared in the early fifties along with the other stop-gap transonic techniques (the wing-flow, the annular tunnel, and the body-drop programs). The final summary of the Langley bump tests of wings by Polhamus (ref. 104) contains the following modest obituary: "There are many shortcomings of the Transonic Bump technique. . . . The results are believed to give at least a qualitative indication of the type of effects encountered at transonic speeds, and fairly reliable indications of trends. . ."